Curvic Seal for Gas Turbine Engine

ABSTRACT

A curvic seal for a gas turbine engine includes a cylindrical seal with a retention feature.

This application is a continuation of U.S. patent application Ser. No.14/774,308, filed Sep. 10, 2015, which is a 371 of PCT/US2014/027002,filed Mar. 14, 2014, which claims benefit of U.S. Provisional PatentAppln. No. 61/784,063 filed Mar. 14, 2013.

BACKGROUND

The present disclosure relates to a gas turbine engine and, moreparticularly, to a coupling mechanism for rotational elements.

A curvic coupling is a self-centering joint that transfers torquebetween rotatable elements. The curvic coupling provides an accurate,light, compact, and self-contained connection in which curvic teethoperate as centering and driving devices. The most widely used type ofcurvic coupling for gas turbine engines is the fixed curvic coupling.Many turboprop, turboshaft and turbofan engines include examples of thisapplication in which disk-like rotatable elements of a rotational groupare mounted together with fixed curvic coupling teeth. Curvic couplings,however, may not be air tight and are often sealed by a curvic seal.

SUMMARY

A curvic seal for a gas turbine engine according to one disclosednon-limiting embodiment of the present disclosure includes a cylindricalseal with an axial retention feature.

In a further embodiment of the present disclosure, the axial retentionfeature is U-shaped.

In a further embodiment of any of the foregoing embodiments of thepresent disclosure, the cylindrical seal is a split-ring sheet metalseal.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes a rolled flange on an aft edge section of thecylindrical seal.

A curvic joint assembly for a gas turbine engine according to anotherdisclosed non-limiting embodiment of the present disclosure includes acurvic coupling between a first component and a second component alongan axis and a curvic seal within an interior of the first component andthe second component with respect to the axis to seal the curviccoupling, the curvic seal includes an axial retention feature thatinterfaces with the first component.

In a further embodiment of any of the foregoing embodiments of thepresent disclosure, the axial retention feature is U-shaped.

In a further embodiment of any of the foregoing embodiments of thepresent disclosure, the axial retention feature interfaces with an innerrim that extends from an inner surface of the first component.

In a further embodiment of any of the foregoing embodiments of thepresent disclosure, the first component and the second component arecylindrical.

In a further embodiment of any of the foregoing embodiments of thepresent disclosure, the curvic seal is a split-ring sheet metal seal.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes a rolled flange on an aft edge section of the curvicseal.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled m the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of one example gas turbine engine;

FIG. 2 is a schematic partial outer view of a curvic coupling;

FIG. 3 is a cross-sectional view of the curvic coupling;

FIG. 4 is a perspective view of a curvic seal for the curvic coupling;and

FIG. 5 is a schematic partial inner view of the curvic coupling.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbo fan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan in the disclosed non-limiting embodiment, it should beappreciated that the concepts described herein are not limited to usewith turbofans as the teachings may be applied to other types of turbineengines such as a turbojets, turboshafts, industrial gas turbines, andthree-spool (plus fan) turbofans wherein an intermediate spool includesan intermediate pressure compressor (“IPC”) between a Low PressureCompressor (“LPC”) and a High Pressure Compressor (“HPC”), and anintermediate pressure turbine (“IPT”) between the high pressure turbine(“HPT”) and the Low pressure Turbine (“LPT”).

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing structures38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 (“LPC”) and a lowpressure turbine 46 (“LPT”). The inner shaft 40 drives the fan 42directly or through a geared architecture 48 to drive the fan 42 at alower speed than the low spool 30. An exemplary reduction transmissionis an epicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A whichis collinear with their longitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed with thefuel and burned in the combustor 56, then expanded over the HPT 54 andthe LPT 46. The turbines 54, 46 rotationally drive the respective lowspool 30 and high spool 32 in response to the expansion. The main engineshafts 40, 50 are supported at a plurality of points by bearingstructures 38 within the static structure 36. It should be appreciatedthat various bearing structures 38 at various locations mayalternatively or additionally be provided.

In one non-limiting example, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 48can include an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3, and in another example is greater thanabout 2.5:1. The geared turbofan enables operation of the low spool 30at higher speeds which can increase the operational efficiency of thelow pressure compressor 44 and low pressure turbine 46 and renderincreased pressure in a fewer number of stages.

A pressure ratio associated with the low pressure turbine 46 is pressuremeasured prior to the inlet of the low pressure turbine 46 as related tothe pressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 44, and the low pressure turbine 46has a pressure ratio that is greater than about five (5:1). It should beappreciated, however, that the above parameters are only exemplary ofone embodiment of a geared architecture engine and that the presentdisclosure is applicable to other gas turbine engines including directdrive turbofans.

In one embodiment, a significant amount of thrust is provided by thebypass flow path due to the high bypass ratio. The fan section 22 of thegas turbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet. Thisflight condition, with the gas turbine engine 20 at its best fuelconsumption, is also known as bucket cruise Thrust Specific FuelConsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“T”/518.7)^(0.5) in which “T” represents the ambienttemperature in degrees Rankine. The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 fps (351 m/s).

With reference to FIG. 2, a curvic coupling 60 is located between afirst component 62 and a second component 64 of an assembly 66 such asin the high spool 32 between the high pressure compressor 52 (“HPC”) andthe high pressure turbine 54 (“HPT”) which may be a blind moduleassembly location. The first component 62 and the second component 64may be defined along axis A.

With reference to FIG. 3, a curvic seal 68 is located over the curviccoupling 60 between the first component 62 and the second component 64.The curvic seal 68, in one disclosed non-limiting embodiment, is asplit-ring sheet metal seal (FIG. 4) that mechanically loads on an innerdiameter of the curvic coupling 60 (also shown in FIG. 5). That is, thecurvic seal 68 is generally cylindrical with a slot 70 (FIG. 4). Themagnitude of the mechanical loading is sufficient to overcome a pressuredrop across the curvic coupling 60 between a high pressure area 72H anda low pressure area 72L.

With continued reference to FIG. 3, the curvic seal 68 includes aretention feature 74 that interfaces with an inner rim 76 on the firstcomponent 62 to provide axial retention. The retention feature 74 in onedisclosed non-limiting embodiment is a U-shaped bend. The curvic seal 68is thereby axially locked to the first component 62 to facilitate blindmodule assembly. A rolled flange 78 on an aft edge section 80 of thecurvic seal 68 further facilitates blind module assembly. That is, therolled flange 78 provides an effective interface as the first component62 is axially assembled to the second component 64 as illustratedschematically by arrow W.

The curvic seal 68 is light-weight, requires no pre-load, is axiallylocked in place and the rolled flange 78 allows for blind moduleassembly such as between the high pressure compressor 52 (“HPC”) and thehigh pressure turbine 54 (“HPT”) without the potential for buckling thecurvic seal 68.

The use of the terms “a” and “an” and “the” and similar references inthe context of description (especially in the context of the followingclaims) are to be construed to cover both the singular and the plural,unless otherwise indicated herein or specifically contradicted bycontext. The modifier “about” used in connection with a quantity isinclusive of the stated value and has the meaning dictated by thecontext (e.g., it includes the degree of error associated withmeasurement of the particular quantity). All ranges disclosed herein areinclusive of the endpoints, and the endpoints are independentlycombinable with each other. It should be appreciated that relativepositional terms such as “forward,” “aft,” “upper,” “lower,” “above,”“below,” and the like are with reference to the normal operationalattitude of the vehicle and should not be considered otherwise limiting.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beappreciated that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

1-10. (canceled)
 11. A curvic seal for a curvic coupling that transferstorque between a first component and a second component along an enginecentral longitudinal axis of a gas turbine engine comprising: acylindrical seal with a retention feature shaped to interface with a rimthat extends from a surface of a cylindrical component such that thecylindrical seal is axially locked to the first component to facilitateblind module assembly of the first component and the second component.12. The curvic seal as recited in claim 11, wherein the retentionfeature is U-shaped.
 13. The curvic seal as recited in claim 11, whereinthe cylindrical seal is a split-ring sheet metal seal.
 14. The curvicseal as recited in claim 11, further comprising a rolled flange on anaft edge section of the cylindrical seal to facilitate the blind moduleassembly.
 15. A curvic seal located along an engine central longitudinalaxis for a curvic coupling that transfers torque between a firstcomponent and a second component of a gas turbine engine, comprising: acylindrical seal with a retention feature shaped to interface with a rimthat extends from a surface of a cylindrical component such that thecylindrical seal is axially locked to the first component to facilitateblind module assembly of the first component and the second component.16. The curvic seal as recited in claim 15, wherein the retentionfeature is U-shaped.
 17. The curvic seal as recited in claim 15, whereinthe cylindrical seal is a split-ring sheet metal seal.
 18. The curvicseal as recited in claim 15, further comprising a rolled flange on anaft edge section of the cylindrical seal to facilitate the blind moduleassembly.
 19. A curvic seal located along an engine central longitudinalaxis to facilitate blind module assembly of a first component and asecond component of a gas turbine engine, comprising: a cylindrical sealwith a retention feature and a rolled flange on an aft edge section ofthe cylindrical seal to facilitate the blind module assembly.
 20. Thecurvic seal as recited in claim 19, wherein the retention featureinterfaces with a rim that extends from a surface of a cylindricalcomponent.
 21. The curvic seal as recited in claim 19, wherein thecylindrical seal is a split-ring sheet metal seal.
 22. A method forblind module assembling a gas turbine engine curvic joint along anengine central longitudinal axis, comprising: attaching a retentionfeature of a cylindrical seal to a rim that extends from a surface of acylindrical component such that the cylindrical seal is axially lockedto a first component; and blind module assembling the first component toa second component along the engine central longitudinal axis to seal acurvic coupling that transfers torque between the first component andthe second component.
 23. The method as recited in claim 22, wherein theretention feature is U-shaped.